Flow deflector for combustion chamber apparatus



0&9, 1956 w. F. EGBERT 2,765,620

FLOW DEFLECTOR FOR COMBUSTION CHAMBER APPARATUS Filed June 23, 1951 2 Sheets-Sheet l ISnventor O 9%7m 'wwf Oct 9, 1956 w. F. EGBERT 2,765,620

FLOW DEFLECTOR FOR COMBUSTION CHAMBER APPARATUS Filed June 25, 1951 2 Sheets-Sheet 2 nnentor KXf FLOW DEFLECTGR FOR COMBUSTION (Ill-IAMBER APPARATUS William F. Egbert, Camby, Ind, assignor to General Motors Corporation, Detroit, Mich, a corporation of Delaware Application June 23, 951, Serial No. 233,171

4 Claims. (Cl. 60-3937) This invention relates to combustion chambers of a type suited for use in gas turbine engines or the like, and is par: ticularly directed to improving the flow or distribution of hot combustion gases discharging from the combustion apparatus into the gas turbine of engines of this type.

In gas turbine engines of the type compnsmg a rotary compressor, a turbine, and combustion apparatusbetween the compressor and turbine, the hot gases discharging from the combustion apparatus are commonly supplied to the turbine through a plurality of individual discharge conduits known as transition liners associated with the combustion apparatus. rdinarily, the engine configuration is such that the transition liner axis is at an angle to the axis of the combustion chamber. Therefore, in passing through the transition liners, the discharging combustion products, usually in the form of a hot gas core, tend to crowd inwardly, for example, thereby creating a non-uniform temperature distribution pattern of flow into the turbine wherein the heat is most intensely concentrated in the inner or stress bearing portion of the turbine blading. Such a temperature distribution is undesirable for, among other reasons, it increases the stresses in the turbine and generally weakens the turbine structure.

Accordingly, it is the principal object of the present invention to improve the flow or distribution of hot combustion gases discharging from the combustion apparatus of gas turbine engines and to obtain a more uniform temperature distribution pattern into the gas turbine.

Another object of the invention is to provide means adapted to be compactly embodied in the combustion apparatus of gas turbine engines for the accomplishment of these general ends without additional space requirements or other modification of the engine.

In accordance with the present invention the hot gases discharging from the combustion apparatus are deflected by suitable flow deflecting means embodied in the transition liners, thereby providing a moreu niform temperature distribution pattern in the gas flowing into the tun bines.

The nature of the present invention together with the features and advantages thereof will be apparent from the following detailed description and drawings of the preferred embodiment thereof, in which:

Fig. 1 is a fragmentary longitudinal sectional view of the discharge end of the combustion apparatus and turbine of a gas turbine engine which includes a preferred embodiment of the invention;

Fig. 2 is a transverse elevation of a part of the structure of Fig. 1 taken substantially in the plane 22 thereof;

Fig. 3 is a fragmentary perspective view of part of the combustion apparatus of Fig. l; and

Fig. 4 is a sectional view of a fragmentary portion of Fig. 1 taken substantially in the plane 4--'i thereof.

The invention is illustrated as incorporated in a known engine, which will be briefly described as an aid to an understanding of the invention. Only the discharge end of one form of combustion apparatus and the turbine portion of an aircraft gas turbine engine is shown and de- 2,765,620 Patented Oct. 9, 1956 scribed herein in the interest of clarity of the drawings and conciseness of the specification since the general structure of such an engine is well understood by those skilled in the art to which this invention pertains.

Referring to the drawings, Fig. 1 is a fragmentary sectional view of the discharge end of one of a plurality of individual combustion chambers 10 the discharge of each of which is supplied to the turbine 12 through individual transition liners 14 contained within a space defined by the outer casing or shroud 15 and an inner casing 16 of the aft frame assembly 18 of the engine.

The combustion chambers 10 are circumferentially arranged and axially supported between the compressor out: let diffuser casing or mid-frame (not shown) and the aft frame 18 forward of the turbine. Each chamber comprises a cylindrical outer casing 20 and a removable inner combustion liner or flame tube 22, the latter having perforations (not shown) about its periphery for purpose of admitting compressor discharge air to support the combustion process taking place therein. The outer combustion chamber casing is double-walled at its aft end as shown, the inner wall 24 thereof being yieldably connected by means of a convoluted bellows seal 26 with a short cylindrical collar or extension duct 28 the diameter of which is less than that of the outer wall of the outer combustion casing 20 so as to permit relative movement therebetween caused by longitudinal expansion of the structure under operating conditions. A corrugated ring 25 acts as a spacer between the wall 24 of the outer duct and the duct extensions 28 thereof. Each flame tube 22 communicates with one of the transition liners 14 through a short transition duct member 30 that extends rearwardly slightly into the inlet of the transition liner with which it is associated.

The transition liners are contained and supported within the aft frame assembly 18, which is a fabricated structure constituted by a rearwardly converging tubular central housing 32, which extends between the engine mid-frame (not shown) and the turbine, and the aft frame inner casing 16 mounted about the aft end of the central housing 32. The aft frame inner casing 16 comprises an annular face plate 34 having a plurality of forwardly extending annular flanges 36 defining a plurality of circular openings therein and a hollow internally reinforced, frustoconical aft section 38 that converges in an aft direction. Reinforcement of the face plate 34 is provided by a plurality of radial support struts 40 positioned between adjacent openings therein and extending obliquely from the rear surface thereof to intersect the conical section 38 of the inner casing. A reinforcing bolting flange ring 42 welded to the rear surface of the face plate 34 has bolted thereto the flanged forward end of the aft frame outer casing 15. Mating flanges 46, 47 are provided about the discharge end of the yieldably connected duct extension 23 of the outer combustion chamber casing 20 and an axially aligned flange 36 of the annular face plate 34, respectively, whereby the discharge end of the individual combustion chambers may be detachably connected to the flanges 36 by means of ring clamps 48 as illustrated.

The turbine 12 comprises a turbine inlet casing, which contains the turbine nozzle 50, and the turbine rotor 52. The turbine nozzle 50 is supported by the aft frame inner casing 16 and comprises a flanged inner ring 54 bolted to the conical section 38 of the inner casing 16 and a stepped outer ring 56 between which is mounted an annular row of radially disposed stationary turbine blades 58 (Fig. 2). The stepped outer ring 56 is supported from the inner ring 54 by blades 58 and forms a support for a spacer ring 60 the opposite surfaces of which are bolted to the flanged aft end of the aft frame outer casing 15 and an exhaust casing 62. The turbine rotor 52 comprises a turbine wheel 64 that may be mounted on a shaft (not shown) passing axially through and mounted on bearings in the tubular aft frame housing 32. Mounted on the outer rim of the turbine wheel 64 are the turbine buckets 66.

The transition liners 14 are each positioned between the discharge end of respective ones of the inner combustion liner flame tubes 22 and the turbine nozzle 50 and structurally present a rearwardly constricting smooth passage the axis of which is inclined to that of the combustion chambers. The inlet ends of the transition liners are supported within respective ones of the flanged openings 36 in the face plate 34 of the aft frame inner casing 16 and are spaced therefrom by a corrugated ring 7%? fixed around the periphery of each liner as shown in Figs. 1 and 2. The inlet end of each transition liner is of circular cross section and smoothly converges in its rearward extension to form a segment of an annulus at its discharge end, the curved inner surface 72 of the liner being inclined to the axis of the combustion chambers and approximately parallel to the surface of the conical section 38 of the aft frame inner casing. Embossments '74 are provided in the outer 76 and inclined inner 72 curved surfaces of the liner to increase the rigidity thereof. The discharge ends of the transition liners are supported by the turbine nozzle 50 and are detachably connected to the stepped outer ring 56 thereof by hinge-like retaining brackets 78, 79 alternately arranged thereon along the joint between the outer surface of the transition liner and the outer ring of the turbine nozzle as illustrated in Fig. 3. Each retainer bracket 7 8, 78 consists of a flat portion welded to the outer surface 76 of the transition liner 14 or the outer ring 56 of the turbine nozzle and ha loops thereon to receive a retaining wire 81 as shown in Fig. 3.

The structure so far described is that of a known aircraft gas turbine engine, and is an example of a suitable environment for the invention. The heated combustion products discharging from the combustion chamber flame tubes 22 are usually in the form of a hot gas core that is most intensely concentrated at the center of the discharge end thereof. In accordance with the invention the hot gas core is deflected outwardly by suitable flow deflecting apparatus installed in the transition liners to provide a more uniform temperature distribution pattern at the discharge end thereof.

The flow deflecting apparatus, in accordance with a preferred embodiment of the invention comprises an inclined deflector plate 86 and a radial support member 88 mounted in the axially constricting region of each transition liner intermediate or approximately midway between the inlet and discharge ends thereof. The deflector plate 86 is inclined outwardly to the axis of the combustion chambers and is in spaced relation with an approximately parallel to the inclined inner surface 72 thereof as illustrated in Fig. l. The deflector plate extends laterally in a transverse direction across the interior of the transition liner as shown in Figs. 2 and 4, the ends of the plate conforming to the diverging side walls of the liner and being bent outwardly and spot welded to the side walls thereof as shown in Fig.4.

The support member $58 may comprise two flanged sections 99, 91, aligned one above the other, that extend radially from the central portion of the deflector plate to which they are welded as shown in Figs. 2 and 4. The flanged ends 99a, 311 of the support member are spot welded to the outer and inner surfaces, respectively, of the transition liner and are reinforced by flanged brackets 93, Q4 similarly welded to the extremities of the radial portion of the support member 38 and to the liner as shown in Fig. 2.

In passing through the axially constricting transition liners, the hot gas core discharging from the combustion chambers will be seen to be deflected outwardly by the deflector plate 86. A more uniform temperature distribution pattern will thus be obtained across the discharge end of the transition liner and in the turbine area.

The plate 86 acts as an airfoil to direct the gas outwardly, minimizing crowding of the gas against the inner surface 72 of the transition liner and the resulting higher gas density, velocity, and temperature at the roots of the turbine blades. Obviously, more elaborate arrangements than the simple plate 86 could be employed, but the arrangement disclosed has the advantage of simplicity and is adequate to obtain the desired result.

Although a specific embodiment of the invention has been shown and described, it will be understood that it is but illustrative and that various modifications may be made therein without departing from the scope and spirit of this invention.

What is claimed is:

1. A gas turbine engine comprising, in combination, a combustion apparatus adapted for flow of gases therethrough in a given direction, a turbine, discharge conduit means generally aligned with and adjacent said combustion apparatus for conducting combustion products discharging therefrom to said turbine, said discharge conduit means having a rearwardly extending walled portion generally parallel to the direction of flow and a rearwardly converging walled portion inclined to the said direction of flow and subjected to the impingement of said combustion products passing therethrough from said combustion apparatus, and flow deflecting means mounted in said discharge conduit means intermediate the ends thereof and having the same general inclination to the said direction of flow as the said inclined walled portion of said discharge conduit means for deflecting said combustion products flowing therethrough away from the said inclined walled portion of aid conduit means, said flow deflecting means extending transversely across the interior of said conduit and being of short length in relation to the length of said conduit.

2. The combination with a transition liner for conveying combustion products from the combustion apparatus of a gas turbine engine to the gas turbine thereof, said transition liner being constituted by a duct having an inlet end, a discharge end, a rearwardly converging walled portion inclined to the direction of flow of said combustion products discharged from said combustion apparatus and a rearwardly extending walled portion generally parallel to the said direction of flow; of a flow deflecting plate extending transversely across the interior of said duct and positioned intermediate the ends thereof, said deflecting plate being of short length in the direction of flow in relation to the rearwardly converging walled portion of said duct and further being spaced slightly from and disposed generally parallel to said converging walled portion and inclined to the other said walled portion of the duct.

3. The combination with a transition liner for conveyeombustion products from the combustion apparatus of a gas turbine engine to the gas turbine thereof, said transition liner being constituted by a duct having an inlet end, a discharge end, a rearwardly converging walled portion inclined to the direction of flow of said combustion products discharged from said combustion apparatus and a rearwardly extending walled portion generally parallel to the said direction of flow; of a flow deflecting plate extending transversely across the interior of said duct and positioned intermediate the ends thereof, and a radial extending support strut bracing the central portion of said deflecting plate and extending between said walled portions of said duct, said deflecting plate being of short length in the direction of flew in relation to the rearwardly converging walled portion of said duct and further being spaced slightly from and disposed generally parallel to said converging wal ed portion and inclined to the other said walled portion of the duct.

4. A combustion apparatus comprising an elongated chamber in which a mixture of fuel and air is burned in passing through said chamber and in which a flame core is formed along the axis of said chamber and is projected from said chamber as the products of combustion are discharged therefrom, a transition conduit extending from the outlet end of said chamber and being formed to pro- References Cited in the file of this patent vide a wall in parallel relation to said axis of said chamber UNITED STATES PATENTS and an opposite wall converging towards said parallel Wall at the outlet end thereof and being disposed obliquely g t g to the axis of said chamber and in the path of the flame 5 2'510572 g: 195(9) core as the core is projected from the chamber, and a 2548886 g g 52 1951 bafile extending transversely across Sald transition conduit 638,745 Nathan H y 1953 in spaced relation to said obliquely disposed Wall and in directly opposed relation to and in the path of said core of said combustion products projected from said chamber 10 along the axis thereof. 

